Thrust control for rocket engines



Dec. 11, 1962 M. J. CORBETT 3,

THRUST CONTROL FOR ROCKET ENGINES Filed Jan. 2, 1959 2 Sheets-Sheet 111, 1962 M. J. CORBETT 3,067,574

THRUST CONTROL FOR ROCKET ENGINES Filed Jan. 2, 1959 2 Sheets-Sheet 2CONTROL A3 m5 Elba- Mamba/l (I Corbe/z 53 zzw flrr s United StatesPatent C) ee son Ramo Wooldrid e Inc. Cleveland h ration of Ohio g r acorpo Filed Jan. 2, 1959, Ser. No. 784,624 2 Claims. (Cl. 6035.6)

This invention relates to apparatus for controlling the thrust of arocket engine of the type having a combustion chamber and an exhaustnozzle through which gases of combustion of a fuel consumed in thecombustion chamber are released to provide thrust for the rocket. Moreparticularly, this invention relates to apparatus for controlling thethrust of such a rocket adapted to burn a gelatinous material asamono-propellant fuel.

In rocket engines of the type adapted to burn a gelatinousmono-propellant fuel which is extruded from a fuel storage tank directlyinto a combustion chamber, there is a tendency forthe back pressuresdeveloped by the combustion of the fuel in the combustion chamber tointerfere with the control of the rate of feed of the fuel into thecombustion chamber. It is desirable that this rate of fuel feed beprecisely controlled as a means of controlling the thrust of the rocket.

it is accordingly an object of the present invention to provideapparatus for controlling the thrust of a rocket engine of the typehaving a combustion chamber and an exhaust nozzle by controlling therate of fuel feed to the combustion chamber independently of any backpressures built up in the combustion chamber.

It is a further object of this invention to provide such controlapparatus particularly adapted for controlling the feed rate of agelatinous mono-propellant fuel to a combustion chamber in response to acomparison of the actual pressure in the combustion chamber with apredetermined desired combustion chamber pressure.

It is a further object of this invention to provide such apparatus, theoperation of which may readily be initiated in response to a remotecommand.

It is a further object of this invention to provide such apparatus whichis light in weight, inherently simple and foolproof, and yet efiicientin operation.

Other objects, features, and advantages of the present invention will bemore fully apparent to those skilled in the art from the followingdetailed description taken in connection with the accompanying drawingsin which like reference characters refer to like parts throughout andwherein:

FIGURE 1 is a diagrammatic partially sectional view of a thrust controlsystem in accordance with a first embodiment of the invention.

FIGURE 2 is a diagrammatic view partly sectional of a rocket enginethrust control system in accordance with a second embodiment of theinvention.

FIGURE 3 is a diagrammatic view partly sectional of a rocket enginethrust control system in accordance with a third embodiment of theinvention.

FIGURE 4 is a diagrammatic view partly sectional of a rocket enginethrust control system in accordance with a fourth embodiment of theinvention.

Rocket engines have, in the past, commonly been fueled by liquid orsolid mono-propellants. More recently, gelatin mono-propellants havebeen developed which combine many of the advantages and eliminates manyof the disadvantages of both solid and liquid propellants. Such gelatinmono-propellant fuels are extrudible plastic viscous slurrys orgelatinous materials. Numerous suitable mono-propellant mixtures can bemade into this form. Such mixtures preferably comprise a stabledispersion of a finely divided insoluble solid oxidizer in a continuousmatrix of an oxidizable liquid fuel.

3,667,574 Patented Dec. 11, 1962 The liquid fuel can be any oxidizableliquid, preferably an organic liquid containing carbon and hydrogen.Such liquid fuels include hydrocarbons such as triethyl, benzine,dodecane and the like; compounds containing oxygen linked to a carbonatom such as esters including methyl maleate, diethyl phthalate, butyloxalate, and the like; alcohols such as benzyl alcohol, triethyleneglycol and the like; others such a methyl o-naphthyl ether and the like,and many others.

The solid oxidizer can be any suitable active oxidizing agent whichyields an oxidizing element such as oxygen, chlorine, or fluorinereadily from combustion of the fuel and which is insoluble in the liquidfuel vehicle. Such oxidizers include inorganic oxidizing salts such asammonia, sodium and potassium per chlorate or nitrate and metalperoxides such as barium peroxide.

Finely divided solid metal powders, such as aluminum or magnesium, maybe incorporated in the mono-propellant composition as an additional fuelcomponent along with the liquid fuel. Such metal powders possess theadvantages both of increasing the fuel density and improving thespecific impulse of the mono-propellant because of their high heats ofcombustion.

Gelling agents for imparting desired cohesiveness and ilowcharacteristics to the plastic mixture include natural and syntheticpolymers such a polyvinyl chloride, polyvinyl acetate, cellulose esterssuch as cellulose acetate, cellulose ethers such as ethyl cellulose,metal salts of higher fatty acids such as the sodium or magnesiumstearates and palmitates.

The amount of oxidizes is prefeably as a stoichiometric level withrespect to the liquid fuel, although minimum concentrations of solidoxidizer as low as 40% by weight are operative. about 65% by weight ofthe mixture. A preferred operative gelatinous mono-propellant includes agel composed of up to 50% by weight of a liquid fuel, from 40% to 65% byweight of an oxidizer and from 3% to 10% of a gelling agent. A specificoperative fuel can be composed of about 56% by weight of solid oxidizerssuch as potassium per chlorate, about 45% by weight of liquid fuels suchas triethyl benzine, and about 5% by weight of a gelling agent such asethyl cellulose. It is to be understood, however, that this invention isnot limited to use with any particular gelatinous mono-propellantmixture, but rather is directed to apparatus for controlling the thrustof a rocket engine of the type adapted to burn such a mono-propellantfuel after is has been extruded into a combustion chamber in order tocontrol the thrust I of the rocket developed by discharge of the gasesof combustion through an exhaust nozzle leading from the combustionchamber.

Turning now to the drawings, there is shown in FIG- URE 1 a partlysectional diagrammatic view of a diagrammatic view of a control systemfor a rocket engine 10 of the type adapted to controllably burn such agelatinous mono-propellant fuel. The engine 10 includes a fuel tank 11in which the gelatinous mono-propellant fuel 12 is stored. A drive meanssuch as the plunger or piston 13 is mounted for sealed slidable motionin the fuel tank 11 in order to extrude the fuel 12 from the outlet end14 of the fuel tank. Mounted on and in open communication with theoutlet end 14 of the fuel tank is a combustion chamber 15 whichpreferably has a plurality of burner tubes 16 therein. Each of theburner tubes 16 also preferably includes a fuel splitting device 17which in practice may be a conical plug mounted on a spider or a Wireextending across the burner tube 16. Gases generated in the combustionchamber 15 are vented through an exhaust nozzle 18 which may beunitarily attached to the combustion chamber. Of

in general, the oxidizer will constitute seamen course, the venting ofthe exhaust gases provides the thrust of the rocket engine.

The plunger 13 ma as shown in FIGURE 1, be driven by an inert gasadmitted under controlled pressure in back of the plunger to urge theplunger toward the outlet end 14 of the fuel tank 11. As shown in FIG- ULE .l, the inert gas used may, for example, be nitrogen stored in a tank13 forming the forward end wall of the fuel tank 11. Tank 19 isconnected through a conduit or line 26 to a pressure regulator 21 andthence .through conduit 22 to enter the fuel tank at a point in back ofthe plunger 13. Interposed in the conduit 26 is a cartridge firedshut-off valve 23 which in storage and ground handling is normally inthe closed position.

Shut-off valve 23 may be controlled by a common electrical circuit whichalso controls an igniter 24 mount ed in the combustion chamber 15.lgniter 24 may be of any well-known type such as an electrically firedsquib or a hot wire variety. Cartridge fired shut-off valve 23isconnected by a wire 25 to a switch 26. Igniter 24 is also connected toswitch 26 by a wire 27. The switch in turn is connected when in theclosed position to a charged capacitor 27 or to any other suitablesource of electrical signal such as a battery or power supply. The otherside of the source 27 is gounded and, of course, the other terminals ofigniter 24 and cartridge fired valve 23 would also be grounded. Whenswitch 26 is closed, an electrical signal is simultaneously applied tothe cartridge fired shut-off valve 23 and to the igniter 24. Thissimultaneously opens the shut-off valve 23 and actuates the igniter 24.Opening of shut-off valve 23 admits nitrogen under pressure in back ofplunger 13 thereby driving it toward the combustion chamber 15 and henceextruding fuel through the burner tubes 16 which is ignited by theigniter 24 to start the opera tion of the rocket engine.

As noted above, the thrust produced by the rocket engine is a functionof the combustion chamber pressure which in turn is a function of theburning rate of the fuel, which in turn is a function of the rate atwhich the fuel is extruded to the combustion chamber. In order tomaintain the thrust at a preselected desired value, a conduit or line 29is connected through an orifice 3b to be in open communication with thecombustion chamber 15. The other end of line 29 is in open communicationwith one side of a thrust regulator actuating chamber 31 having acentrally positioned diaphragm 32 therein. An adjustable calibratedspring 33 urges the diaphragm 32 to the left as seen in FIGURE 1,whereas the combustion chamber gas pressure as applied through line 29to the chamber 31 urges the diaphragm 32 to the right as seen inFIGURE 1. Mounted on diaphragm 32 is a valve stem 34 and a valve 35which may seat on a valve seat 36 at the inlet to the primary chamber3'7 of the thrust regulator 21'. A second diaphragm member 38 ispositioned to divide the chamber 37 into two portions. A spring 39 isconnected between valve 35 and one side of the diaphragm 38 to transmitthe resultant'force from the spring 33 and the pressure in thecombustion chamber introduced through conduit 29, to the diaphragm 38,the characteristics of the spring 39 complementing the characteristicsof the spring 33 to afford a desired force transmitting capability forthis purpose. A valve stem 4ft carrying a valve 41 is connected to theother side of the diaphragm 38. Valve 41 is positioned to cooperate witha valve seat 42 at the inlet to chamber 37 from the nitrogen line 2% atthe point where it enters the thrust regulator 21. The spring 39 also iseffective in transmitting the incremental force from the nitrogen line29 which bears against the diaphragm 38, to the valve 35 and thence tothe diaphragm 32. Accordingly, the position of valve 41 will be theresultant of the pressure of the gas bearing against the diaphragm 58,and the difierential forcefromthe pressure .of the gas bearing againstthe diaphragm 32 (and admitted by valve 35 to bear against diaphragm 38)and force afforded by the spring 33.

In operation, the adjustable calibrated spring 33 is set for a desiredvalue of thrust as by adjustment of the position of a plunger 33abearing against spring 33. This .loads the spring 33, diaphragm 32,valve stem 34, spring 39, and diaphragm 38 and valve stem 40 to open thenitrogen valve 41 so as to admit nitrogen under pressure through line 22to actuate the drive piston 13 thereby causing the extrusion of fuel 12into the combustion chamber 15. Burning of the fuel in the combustionchamber builds up pressure therein which is comunicated through line 29to act on the diaphragm 32 and urge it toward the right in FIGURE 1thereby tending to urge the valve 41 towards its closed position toreduce the pressure of the nitrogen being supplied through line 22 andthereby reduce the rate of fuel extrusion; When the desired value ofrocket thrust has been achieved, as indicated by the desired value ofcombustion chamber pressure, the pressure acting on the left side ofdiaphragm 32 is equalto the pressure applied to the right side of thediaphragm by spring 33 and the diaphragm 32 is in equilibrium. In thisequilibrium position, the valve 41 is positioned to maintain a nitrogenpressure sufiicient to produce a rate of fuel extrusion which will justmaintain the desired thrust. However, it will be apparent that thepressure in nitrogen tank 19 itself falls as more and more nitrogen isused during the course of the rockets travel. By way of ex ample, theinitial pressure of the nitrogen may be 2,000 p.s.i and the finalpressure at the end of expulsion may be as low as 400' p.s.i. Suchpressures may be used, for example, where it is desired to maintain acombustion chamber pressure in the neighborhood of 300 psi It isapparent that as long as fuel extrusion is desired to continue, thenitrogen pressure must be greater than the combustion chamber pressurein order to overcome the back pressure from the combustion chamber onthe fuel.

It will, of course, be understood that the thrust regulator calibratedspring 33 may either be permanently adjusted by manual means beforefiring the rocket in order to maintain a fixed predetermined value ofthrust, or that any suitable remotely controlled actuating means may beprovided to .actuate a plunger 33a so as to vary the ad justment of thespring 33 to produce a programmed variation of thrust during the flightof the rocket. If, for example, it is desired to entirely stopcombustion or reduce thrust to zero, it is only necessary to reduce thesetting of the thrust regulator to a value less than the minimum valueof combustion chamber pressure necessary to support combustion of thefuel. In this respect, it will be understood that most mono-propellantfuels of the type described above, have a relatively high burning rateand that this total burning rate is in part determined by the totalsurface area of the fuel exposed for burning by the fuel splitters 17which form the conical surface of the fuel in the burner tubes 16.- Ifthe rate of fuel extrusion is reduced to the point where the conicalsurfaces are burned away so that only the fiat circular cross-sectionalsurface of fuel in each burner tube remains, combustion chamber pressureis reduced to a point less than that necessary to support burning of thefuel. A typical value of such necessary minimum combustion chamberpressure for many such fuels is approximately 200 p.s.i. Of course, ifdesired, this reduction in combustion chamber pressure could also beachieved upon a remote command by providing a blow-out valve or otherpressure relief means in thecombustion chamber downstream of the burnertubes. Such additional means is, however, not in general necessary.

For any given setting of the adjustable thrust regulator spring 33,either a fixed or programmed variable setting, the control system shownin FIGURE 1 acts a servo-system to maintain the combustion chamberpressure at the desired value independently of changes in the nitrogensupply pressure from tank 19 during the course of operation. Thus,assuming that immediately after the cartridge fired shut-off valve 23 isopened, the system establishes the pretermined desired combustionchamber pressure in accordance with the mode of operation discussedabove for the initial value of nitrogen pressure. Then, as the nitrogenpressure begins to decrease, the rate of fuel extrusion will also tendto decrease. This in turn tends to decrease the combustion chamberpressure thereby permitting the diaphragm 32 to move to the left as seenin FIGURE 2 and thus tending to further open the valve 41 so as tocompensate for the fall in pressure in tank 19 and thereby increase thepressure on the plunger 13 to its original value. Of course, it will beunderstood that the original value of the pressure on plunger 13 is lessthan either the initial or final value of pressure in tank 19 but isgreater than the desired pressure in combustion chamber in order toovercome the back pressure therefrom. The desired differences in valuebetween the nitrogen pressure acting on the plunger 13 and the nitrogenpressure existing in tank 19 is maintained by varying the pressure dropthrough valve 41 in the manner described above.

In FIGURE 2, there is shown a second embodiment of the control systemwherein a solid grain gas expulsion system is used in place of thestored nitrogen or other inert gas. In other respects, the rocket is thesame as that shown in FIGURE 1 and like parts are indicated by likereference characters and will not be further described. The over-allconstruction for the solid grain gas expulsion system is lighter inweight than that required for the nitrogen system. This arises from thefact that the nitrogren tank 19 tends to be a heavy and cumbersomeelement. In the system of FIGURE 2, therefore, gas or fluid foroperating the piston 13 is generated from a low temperature solidpropellant 45, which may, for example, be ammonium nitrate fuel or gunpowder, stored in a chamber 46 and adapted to be originally ingnited byan igniter 47 in the same manner that the cartridge fired shut-oif valve23 is initially opened-in the system of FIGURE 1. The solid grain gasesgenerated by burning of the fuel in chamber 47 are preferably cooled bythe evaporation of a water-alcohol mixture contained within refractoryfibers 48 packed into the end of the grain chamber 46. At 350 p.s.i.,for example, a 60-40 mixture of alcohol and Water (freezing point 70 F.)will evaporate at 400 F. and absorb about 600 B.t.u.s per pound ofdiluent. Since a low flames temperature such as 1800" F. of solidpropellant will release 681 B.t.u. per pound in dropping to 400 F.,approximately 50% of the gas flow for expulsion is supplied by theexaporated diluent.

The solid grain gases (diluted to 400 F.) are admitted behind theexpulsion piston 13 through a double ported regulator valve 49 connectedto be operated by the valve stem 34 from the diaphragm 32 of thrustregulator chamber 31. The valve 49 is designed so that the sum of theareas of the two ports 50 and 51 is constant independently of theposition of the reciprocating double valve 52. This constant areaisprovided in order to minimize possible effects on expulsion graincombustion. It will be noted. that the solid fuel chamber 46 isconnected by av conduit or pipe line 53 to a central point between thetwo ports 50 and 51 of the double ported valve 49. Depending upon theposition of the reciprocable double valve 52, gases from line 53 willeither be exhausted to the'a'mbient atmosphere through a conduit 54which connects with port 51, or will be supplied through a conduit 55which connects with port 50 to the forward side of the plunger 13. Ofcourse, the gases from conduit 53 may divide and flow partly througheach of the conduits 54 and 55 respectively, in proportions which aredetermined by the relative position of the double valve 52 with respectto the ports 50 and 51 as will be 6 obvious to those skilled in the art.The position of the double valve 52 is in turn determined by theposition of the valve stem 34 attached to diaphragm 32 since valve 52 isintegrally atached to valve stem 34. It is thus apparent that the doublespool valve 52 acts as a flow divider for the gases generated from thesolid fuel in chamber 46. These gases must either leave through exhaustline 54 or be fed to the piston 13 through line 55. Since the spoolvalve is so proportioned that the sum of the areas of the orifices 50and 51 isconstant for any position of the valve 52, the pressure inchamber 48 is maintained at a constant value which is of course thevalue of pressure normally generated as described above. It follows thatthe chamber 46 does not need to be of heavy construction in order toaccomodate an initially high storage pressure as in the inert gasactuated system such as shown in FIGURE 1. And the portion of thechamber 31 receiving the stem 34 may be formed to substantiallyeliminate or minimize flow from the conduit 29 therethrough so as tomaintain the described pressure from the combustion chamber against thediaphragm 32. Thus the pressure from the conduit 29 will not, in thisinstance, directly act on the valve 52. In all other respects, thesystem of FIGURE 2 operates in the same manner as that of FIGURE 1.

Turning now to FIGURE 3, there is shown a similar rocket in which theplunger 13 is drawn toward the com bustion chamber to extrude fuel 12 bymeans of a hydraulically actuated pulled and winch system. It will benoted that the fuel tank 11 as shown in FIGURE 3 has provided therein apulley 60 which is mounted on an axle extending transversely of theinterior of the tank 11. The fuel tank 11 is also provided with a sealedpulley belt outlet 61 through which the belt 62 enters the fuel tank.Belt 62 is attached to plunger 13 on the forward side thereof and istrained over the pulley 60 thence through the sealed opening 61 in theside of the tank to extend in operative wound relationship around awinch 63 Winch 63 is driven by a hydraulic motor 64 which may be anytype of motor wherein the speed of rotation of the output shaft isproportional to the pressure of hydraulic fluid applied through inputline 22 to drive the hydraulic motor 64. Hydraulic motor 64 may, forexample, be of the type shown in US. Patent No. 2,- 779,296, issued onJanuary 29, 1957, to Edward C. Dudley.

The hydraulic motor 64 is connected by a belt or any other suitablemechanical drive means 65 to drive the winch 63 at a speed proportionalto the speed of rotation of the output shaft of the motor. As notedabove, hydraulic fluid for operating the motor 64 is supplied through aninput line 22 and a return line 22a from a control unit 21. The controlunit 21 may be structurally the same as the control unit 21 shown forthe nitrogen supply system of FIGURE 1. The control unit is thereforeshown in FIGURE 3 only in block form.

It will be noted in particular that hydraulic fluid under suitablepressure may be supplied from the missile hydraulic supply system or anyother convenient source through an input line 20 leading to the controlunit 21. It will, of course, be understood that as in the system ofFIGURE 1, the line 20 will lead to a control valve such as the controlvalve 41 shown in FIGURE 1. The control valve is regulated as to thepressure drop thereacross in response to the pressure existing incombustion chamber 15 by means of comparison directly with this pressurecommunicated through line 29 to the control system 21 as in FIGURE 1.The same type of spring biased bellows arrangement would be used tocontrol the valve controlling the pressure drop of the hydraulic fluidthrough line 20. The output line 22 from the control unit in the systemof FIGURE 3, instead of applying nitrogen gas directly to the plunger asin FIGURE 1, applies the hydraulic fluid of controlled pressure to thehydraulic motor 64. This fluid after operating the motor is returnedthrough a hydraulic fluid return line 22a back to the missile supplysource.

It will be apparent that the system of FIGURE 3 operates on principlesentirely analyogous to that of the system of FIGURE 1. The control ofthe pressure of the hydraulic fluid through lines 20 and 22 in responseto combustion chamber pressure in turn controls the speed of operationof the hydraulic motor driving the winch and therefore determining thespeed with which the plunger 13 is pulled toward the combustion chamberby belt 62 riding on the pulley 60. This, of course, in turn controlsthe burning rate of the fuel and hence the combustion chamber pressure.In applications where the missile is already equipped with a suitablesource of hydraulic pressurized fluid, the system shown in FIG- URE 3 issomewhat more economical and simplier than the systems shown in FIGURES1 and 2.

In FIGURE 4 there is shown another hydraulically operated system whereinthe plunger 13 in fuel tank 11 is directly driven by a screw jack 70which in turn is driven by a shaft 71 connected to hydraulic motor 64.The screw jack 70 may be centrally mounted to extend longitudinally oftank 11 between any suitable journaling means 72 at the one end of thetank and a journaling means 73 supported by a spider arrangement 74 inthe outlet of the fuel tank leading to the combustion chamber 15. A bootseal 75 may be provided around the screw jack 70 between the plunger 13and the journaling means 73 in order to isolate the fuel 12 from theother side of the plunger.

Asin the arrangement shown in FIGURE 3, the hydraulic motor 64 may beoperated from any convenient source of pressurized hydraulic fluid suchas an auxiliary power unit or the missiles hydraulic fluid supply so as,to t apply hydraulic fluid through an input line 20 leading to controlunit 21 which may be the same as that shown and described in connectionwith FIGURE 3. The control unit 21 regulates the speed of motor 64 in aservo system responsive to the pressure in the combustion chamberthrough line 29. The hydraulic fluid, after leaving the motor, isreturned to the missile system through return line 221;. Since theprinciples of operation of the system of FIGURE 4 are essentially thesame as those discussed for the other embodiments above, it is notbelieved necessary to discuss the operation in further detail.

Any one of the embodiments shown above provides a dependable system forcontrolling either to a fixed level or to a modulated level, the thrustof a rocket engine adapted to burn a gelatinous mono-propellant fuelextruded into a combustion chamber. Each of the systems affords zerosensitivity to initial propellant temperature and allow for thepossibility of thrust modulation for the well-known boost-coasttechnique; also each of the systems affords an eflicient servo controlof rocket thrust responsive to combustion chamber pressure to accelerateor dccelerate gel fuel flow as required to reach a predeterminedequilibrium operating condition. Each of the systems is such that it isnot impeded by back pressures developed in the combustion chambers andwill afford stable operating performance. The selection between thevarious systems shown is, of course, based on the particular type ofapplication for which the rocket engine is intended. A

While a particular exemplary preferred embodiment of the invention hasbeen described in detail above, it will be understood that modificationsand variations therein may be effected without departing from the truespirit and scope of the novel concepts of the present invention asdefined by the following claims.

I claim as my invention:

1.. Apparatus for controlling the thrust of a gelatinous mono-propellantfueled rocket engine of the type having a combustion chamber and anexhaust nozzle comprising, a fuel tank adapted to contain saidgelatinous mono-propellant fuel, said fuel tank communicating at one endwith said combustion chamber, plunger drive means slidably mounted insealed relation with said fuel tank for axialmovement therein to extrudesaid gelatinous fuel to said combustion chamber, solid fuel means forgenerating as gas under pressure, means to supply said pressurized gasto the other end of said fuel tank to actuate said drive means, adiaphragm biased one one side by calibrated spring means to establish apreselected desired value of thrust of said rocket, means connectingsaid combustion chamber to the other side of said diaphargm to compareactual combustion chamber pressure with said preselected desired value,and valve means positioned to control the flow of said pressurized gasto said drive means, said valve means being connected to be actuated bysaid diaphragm to be responsive to the difference between saidpreselected desired combustion chamber pressure and the actualcombustion chamber pressure to control the pressure of the gas suppliedto said drive means and hence the rate at which said drive means feedssaid fuel to said combustion chamber so as to reduce said difference tozero to maintain said rocket thrust at said preselected value.

2. Apparatus for controlling the thrust of a gelatinous mono-propellantfueled rocket engine of the type having a combustionchamber and anexhaust nozzle comprising, a fuel tank adapted to contain saidgelatinous monopropellant fuel, said fuel tank communicating at one endwith said combustion chamber, plunger drive means slidably mounted insealed relation with said fuel tank for axial movement therein toextrude said gelatinous fuel to said combustion chamber, solid fuel gasgenerating means, atwo seat spool valve, means connected to supply saidsolid fuel gas through one side of said spool valve to actuate saiddrive means, the other side of said spool valve being open to exhaust,the two seat orifices of said spool valve having a constant total areafor any position of said valve so that the pressure in said solid fuelgas generator. is maintained constant for any position for said valve,calibrated means to establish a preselected desired value of combustionchamber pressure corresponding to'a desired value of thrust of saidrocket, means to compare actual combustion chamber pressure with saidpreselected desired value, and means connected to position said spoolvalve responsively to the difference between said preselected desiredvalue of combustion chamber pressure and said actual value of combustionchamber pressure to control the pressure of the gas supplied to saidplunger drive means and hence the rate at which said drive means feedssaid fuel to said combustion chamber so as to reduce said difference tozero to maintain said rocket thrust at said preselected value.

References Cited in the file of this patent UNITED STATES PATENTS2,703,960 Prentiss Mar. 15, 1955 2,816,419 Mueller Dec. 17, 1957 FOREIGNPATENTS 582,621 Great Britain Nov. 22, 1946 OTHER REFERENCESAstronautics, February 1958 (pages 34 and 35 relied on).

Rocket Propulsion Elements, Sutton 2nd edition 1956, published by JohnWiley & Sons Inc. (pages 297-299 relied on).

